1. Field of the Invention
The present invention relates generally to gas turbine engines, and, more specifically, to a turbine support assembly therein, and to gas turbine engine rotor attachment and particularly to heat shields and axial retention for bolts used to connect turbine shafts.
2. Description of Related Art
In an exemplary conventional turbofan gas turbine engine, a compressor is driven by a high pressure turbine through a drive shaft extending therebetween. And, a fan is driven by a low pressure turbine through a fan shaft extending therebetween which is disposed coaxially with the high pressure turbine drive shaft. The fan shaft is typically supported to a stationary casing by frames such as a fan frame and turbine rear frame having roller bearings which support radial loads transmitted by the fan shaft. The low pressure turbine must be suitably supported for preventing unacceptably large clearances between the blade tips thereof and the casing surrounding the blade tips for reducing performance losses due to leakage of combustion gases over the blade tips. Both axial and radial applied forces acting on the low pressure turbine must be suitably accommodated to prevent undesirable variations in blade tip clearances.
More specifically, the low pressure turbine may conventionally include an individual disk from which the rotor blades extend, with the disk being removably fixedly joined to the fan shaft by bolts for example. In another embodiment, the low pressure turbine can include several stages with respective disks and blades extending therefrom, with the disks being suitably joined together by being bolted or welded. The several stages may be collectively joined to the fan shaft by either two axially spaced apart support shafts, or a single support shaft typically disposed near the middle of the low pressure turbine. The two support shaft embodiment distributes the turbine loads to the fan shaft through two axially spaced apart planes and, therefore, more effectively supports the low pressure turbine for reducing variations in the blade tip clearances. However, the single support shaft embodiment is inherently more flexible than the two support shaft embodiment which allows the applied forces to develop bending moments around the single support shaft which elastically deflect the low pressure turbine and, therefore, vary the blade tip clearances of the several stages. Due to the flexibility of the single support shaft, the blade tip clearances of the several stages can vary axially from stage to stage as well as circumferentially around each stage.
Furthermore, in one arrangement, the single support shaft can be removably bolted to the fan shaft, with the fan shaft being supported by bearings at two planes or more. During manufacture of the engine, the low pressure turbine module which is initially separated from the fan shaft is typically balanced as a module before assembly to the fan shaft. Since the support shaft is bolted to the fan shaft upon assembly, and does not otherwise have a separate bearing support, a balancing arbor must be used to simulate the aft end of the fan shaft and its bearings to support the low pressure turbine in the balancing machine. Once balanced, the low pressure turbine may be bolted to the fan shaft and its bearings and supported thereby.
Similarly, in this exemplary embodiment, during maintenance of the engine requiring the removal of the low pressure turbine, another arbor is required to support the low pressure turbine if the fan shaft is removed therefrom.
Gas turbine engines conventionally transfer rotational mechanical energy from turbine sections of the engine to the fan and compressor sections via shaft assemblies that are bolted together for easy assembly and disassembly. Ease of assembly and disassembly provides many benefits from a cost standpoint and enhance modular designs which have been developed to help ship and install gas turbine engines. Examples of such modular engines are disclosed in U.S. Pat. No. 3,842,595 entitled "Modular Gas Turbine Engine" by Smith et al and in U.S. Pat. No. 3,823,553 entitled "Gas Turbine With Selfcontained Power Turbine Module" by Smith, both assigned to the same assignee as the present invention.
Due to the large size of modern high bypass ratio fanjet engines, and particularly their fan sections, it has become useful to incorporate modularity into engine designs. Modularity enhances the engine's assembly and disassembly and facilitates shipment of the engine and its parts for original installation, overhaul, repairs, and retrofitting.
Bolt assemblies to secure shaft sections for the low pressure turbine are often not accessible from the rear during assembly. Axial bolt retention means are required during assembly because such bolt heads lie in closed cavities. Heat shielding the bolts from the hot gases passing through rear stages of the low pressure turbine is also highly desirable if not often required to prolong the useful life of the assembly and increase engine reliability.
Prior engine designs such as the one illustrated in FIG. 1 have incorporated split ring retainers 2 disposed in circumferential grooves 4 formed in the shank 6 of and to retain the bolt 8 which is used to attach a forward rotor element 10 to an aft rotor element 12. Another prior art design is illustrated in FIG. 2, and uses an individual bolt hook 14 to retain the bolt 8 which attaches forward rotor element 5 to aft rotor element 7. Yet another apparatus shown in the prior art is a retention clip disclosed in U.S. Pat. No. 4,887,949 entitled "Bolt Retention Apparatus" by Dimmick, III et al. All of these designs add weight, increase the number of engine parts, and increase the complexity of the engine and assembly and disassembly procedures. Circumferential grooves reduce the shanks load carrying capability and split rings and clips are subject to and may also introduce undesirable engine vibrations.